Nozzle film cooling with alternating compound angles

ABSTRACT

A nozzle segment for a nozzle ring of a gas turbine engine is disclosed. The nozzle segment includes a first endwall, a second endwall, and an airfoil extending between the first endwall and the second endwall. The airfoil includes a multiple groups of cooling apertures spaced apart and alternating in directionality such that a first grouping of cooling apertures is angled toward the first endwall, a second grouping of cooling apertures is angled toward the second endwall and spaced apart from the first grouping of cooling apertures, and a third grouping of cooling apertures are angled toward the first endwall and spaced apart from the second grouping of cooling apertures.

TECHNICAL FIELD

The present disclosure generally pertains to gas turbine engines, and ismore particularly directed toward nozzle segments including film coolingholes with alternating compound angles.

BACKGROUND

Gas turbine engines include compressor, combustor, and turbine sections.The turbine section is subject to high temperatures. In particular, thefirst stages of the turbine section are subject to such hightemperatures that the first stages are often cooled with air directedfrom the compressor and into, inter alia, the nozzle segments andturbine blades.

A portion of the air directed into the nozzle segments may be directedthrough the walls of the nozzle segment airfoils and along the pressureside surface of the walls to film cool the walls. U.S. Pat. No.7,377,743 to D. Flodman discloses a turbine nozzle that includes a midvane mounted between a pair of end vanes in outer and inner bands. Themid vane includes a first pattern of film cooling holes configured todischarge more cooling air than each of the two end vanes havingrespective second patterns of film cooling holes.

The present disclosure is directed toward overcoming one or more of theproblems discovered by the inventors or that is known in the art.

SUMMARY OF THE DISCLOSURE

A nozzle segment for a nozzle ring of a gas turbine engine is disclosed.The nozzle segment includes a first endwall, a second endwall, and anairfoil extending between the first endwall and the second endwall. Theairfoil includes a leading edge, a trailing edge, a pressure side wall,and a suction side wall. The leading edge extends radially from thefirst endwall to the second endwall. The trailing edge extends radiallyfrom the first endwall to the second endwall axially distal to theleading edge. The pressure side wall extends from the leading edge tothe trailing edge. The suction side wall also extends from the leadingedge to the trailing edge. The airfoil also includes a plurality ofshowerhead cooling apertures, a plurality of forward cooling apertures,and a plurality of intermediate cooling apertures. The plurality ofshowerhead cooling apertures span along the leading edge. The pluralityof forward cooling apertures are grouped together proximate theplurality of showerhead cooling apertures. The plurality of intermediatecooling apertures are grouped together in the pressure side wall betweenthe trailing edge and the plurality of forward cooling apertures. Theplurality of showerhead cooling apertures, the plurality of forwardcooling apertures, and the plurality of intermediate cooling aperturesalternate in directionality such that the plurality of showerheadcooling apertures are angled toward the first endwall, the plurality offorward cooling apertures are angled toward the second endwall, and theplurality of intermediate cooling apertures are angled toward the firstendwall.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic illustration of an exemplary gas turbine engine.

FIG. 2 is a perspective view of a nozzle segment for the gas turbineengine of FIG. 1.

FIG. 3 is a cross-sectional view of a portion of the nozzle segment ofFIG. 2 showing the showerhead cooling apertures.

FIG. 4 is a detailed view of the forward cooling apertures of FIG. 2.

FIG. 5 is a detailed view of the intermediate cooling apertures of FIG.2.

FIG. 6 is a detailed view of the aft cooling apertures of FIG. 2.

FIG. 7 is a cross-section of the airfoil of FIG. 2.

DETAILED DESCRIPTION

The systems and methods disclosed herein include a nozzle segment for anozzle ring of a gas turbine engine. In embodiments, the nozzle segmentincludes an upper endwall, an inner endwall, and one or more airfoilsthere between. Each airfoil includes spaced apart groups of coolingapertures through the pressure side wall of the airfoil. One group isangled toward the lower endwall and the next group is angled towards theupper endwall in an alternating pattern for subsequent groups of coolingholes. Alternating the directionality of the groups of cooling aperturestowards the lower endwall and the upper endwall may reduce thetemperatures of the lower endwall and the upper endwall, and may reducethe amount of cooling air needed to effectively cool the nozzle segment.

FIG. 1 is a schematic illustration of an exemplary gas turbine engine100. Some of the surfaces have been left out or exaggerated (here and inother figures) for clarity and ease of explanation. Also, the disclosuremay reference a forward and an aft direction. Generally, all referencesto “forward” and “aft” are associated with the flow direction of primaryair (i.e., air used in the combustion process), unless specifiedotherwise. For example, forward is “upstream” relative to primary airflow, and aft is “downstream” relative to primary air flow.

In addition, the disclosure may generally reference a center axis 95 ofrotation of the gas turbine engine, which may be generally defined bythe longitudinal axis of its shaft 120 (supported by a plurality ofbearing assemblies 150). The center axis 95 may be common to or sharedwith various other engine concentric components. All references toradial, axial, and circumferential directions and measures refer tocenter axis 95, unless specified otherwise, and terms such as “inner”and “outer” generally indicate a lesser or greater radial distance from,wherein a radial 96 may be in any direction perpendicular and radiatingoutward from center axis 95.

A gas turbine engine 100 includes an inlet 110, a shaft 120, acompressor 200, a combustor 300, a turbine 400, an exhaust 500, and apower output coupling 600. The gas turbine engine 100 may have a singleshaft or a dual shaft configuration.

The compressor 200 includes a compressor rotor assembly 210, compressorstationary vanes (stators) 250, and inlet guide vanes 255. Thecompressor rotor assembly 210 mechanically couples to shaft 120. Asillustrated, the compressor rotor assembly 210 is an axial flow rotorassembly. The compressor rotor assembly 210 includes one or morecompressor disk assemblies 220. Each compressor disk assembly 220includes a compressor rotor disk that is circumferentially populatedwith compressor rotor blades. Stators 250 axially follow each of thecompressor disk assemblies 220. Each compressor disk assembly 220 pairedwith the adjacent stators 250 that follow the compressor disk assembly220 is considered a compressor stage. Compressor 200 includes multiplecompressor stages. Inlet guide vanes 255 axially precede the compressorstages.

The combustor 300 includes one or more fuel injectors 310 and includesone or more combustion chambers 390.

The turbine 400 includes a turbine rotor assembly 410 and turbinenozzles 450. The turbine rotor assembly 410 mechanically couples to theshaft 120. As illustrated, the turbine rotor assembly 410 is an axialflow rotor assembly. The turbine rotor assembly 410 includes one or moreturbine disk assemblies 420. Each turbine disk assembly 420 includes aturbine disk that is circumferentially populated with turbine blades. Aturbine nozzle 450 or nozzle ring axially precedes each of the turbinedisk assemblies 420. Each turbine nozzle 450 includes multiple nozzlesegments 451 grouped together to form a ring. Each turbine disk assembly420 paired with the adjacent turbine nozzle 450 that precede the turbinedisk assembly 420 is considered a turbine stage. Turbine 400 includesmultiple turbine stages.

The turbine 400 may also include a turbine housing 430 and turbinediaphragms 440. Turbine housing 430 may be located radially outward fromturbine rotor assembly 410 and turbine nozzles 450. Turbine housing 430may include one or more cylindrical shapes. Each nozzle segment 451 maybe configured to attach, couple to, or hang from turbine housing 430.Each turbine diaphragm 440 may axially precede each turbine diskassembly 420 and may be adjacent a turbine disk. Each turbine diaphragm440 may also be located radially inward from a turbine nozzle 450. Eachnozzle segment 451 may also be configured to attach or couple to aturbine diaphragm 440.

The exhaust 500 includes an exhaust diffuser 520 and an exhaustcollector 550. The power output coupling 600 may be located at an end ofshaft 120.

FIG. 2 is a perspective view of a nozzle segment 451 for the gas turbineengine 100 of FIG. 1. Nozzle segment 451 includes upper shroud 452,lower shroud 456, airfoil 460, and second airfoil 470. In otherembodiments, nozzle segment 451 can include more or fewer airfoils.Upper shroud 452 may be located adjacent and radially inward fromturbine housing 430 when nozzle segment 451 is installed in gas turbineengine 100. Upper shroud 452 includes upper endwall 453. Upper endwall453 may be a portion or a sector of an annular shape, such as a sectorof a toroid or a sector of a hollow cylinder. The toroidal shape may bedefined by a cross-section with an inner edge including a convex shape.Multiple upper endwalls 453 are arranged to form the annular shape andto define the radially outer surface of the flow path through a turbinenozzle 450. Upper endwall 453 may be coaxial to center axis 95 wheninstalled in the gas turbine engine 100.

Upper shroud 452 may also include upper forward rail 454 and upper aftrail 455. Upper forward rail 454 extends radially outward from upperendwall 453. In the embodiment illustrated in FIG. 2, upper forward rail454 extends from upper endwall 453 at an axial end of upper endwall 453.In other embodiments, upper forward rail 454 extends from upper endwall453 near or adjacent to an axial end of upper endwall 453. Upper forwardrail 454 may include a lip, protrusion or other features that may beused to secure nozzle segment 451 to turbine housing 430.

Upper aft rail 455 may also extend radially outward from upper endwall453. In the embodiment illustrated in FIG. 2, upper aft rail 455 is ‘L’shaped, with a first portion extending radially outward from the axialend of upper endwall 453 opposite the location of upper forward rail454, and a second portion extending in the direction opposite thelocation of upper forward rail 454 extending axially beyond upperendwall 453. In other embodiments, upper aft rail 455 includes othershapes and may be located near or adjacent to the axial end of upperendwall 453 opposite the location of upper forward rail 454. Upper aftrail 455 may also include other features that may be used to securenozzle segment 451 to turbine housing 430.

Lower shroud 456 is located radially inward from upper shroud 452. Lowershroud 456 may also be located adjacent and radially outward fromturbine diaphragm 440 when nozzle segment 451 is installed in gasturbine engine 100. Lower shroud 456 includes lower endwall 457. Lowerendwall 457 may be a portion or a sector of an annular shape, such as atoroid or a hollow cylinder. The toroidal shape may be defined by across-section with an outer edge including a convex shape. Multiplelower endwalls 457 are arranged to form the annular shape and to definethe radially inner surface of the flow path through a turbine nozzle450. Lower endwall 457 may be coaxial to upper endwall 453 and centeraxis 95 when installed in the gas turbine engine 100.

Lower shroud 456 may also include lower forward rail 458 and lower aftrail 459. Lower forward rail 458 extends radially inward from lowerendwall 457. In the embodiment illustrated in FIG. 2, lower forward rail458 extends from lower endwall 457 at an axial end of lower endwall 457.In other embodiments, lower forward rail 458 extends from lower endwall457 near or adjacent to an axial end of lower endwall 457. Lower forwardrail 458 may include a lip, protrusion or other features that may beused to secure nozzle segment 451 to turbine diaphragm 440.

Lower aft rail 459 may also extend radially inward from lower endwall457. In the embodiment illustrated in FIG. 2, lower aft rail 459 extendsfrom lower endwall 457 near or adjacent to the axial end of lowerendwall 457 opposite the location of lower forward rail 458. In otherembodiments, lower aft rail 459 extends from the axial end of lowerendwall 457 opposite the location of lower forward rail 458. Lower aftrail 459 may also include a lip, protrusion or other features that maybe used to secure nozzle segment 451 to turbine diaphragm 440.

Airfoil 460 extends between upper endwall 453 and lower endwall 457.Airfoil 460 includes leading edge 461, trailing edge 462, pressure sidewall 463, and suction side wall 464. Leading edge 461 extends from upperendwall 453 adjacent an axial end of upper endwall 453 to lower endwall457 adjacent an axial end of lower endwall 457. Leading edge 461 may belocated near upper forward rail 454 and lower forward rail 458. Trailingedge 462 extends from upper endwall 453 distal to leading edge 461,adjacent the axial end of upper endwall 453 opposite the location ofleading edge 461 and from lower endwall 457 adjacent the axial end ofupper endwall 453 opposite or axial distal to the location of leadingedge 461. When nozzle segment 451 is installed in gas turbine engine100, leading edge 461, upper forward rail 454, and lower forward rail458 may be located axially forward and upstream of trailing edge 462,upper aft rail 455, and lower aft rail 459. Leading edge 461 may be thepoint at the upstream end of airfoil 460 with the maximum curvature andtrailing edge 462 may be the point at the downstream end of airfoil 460with maximum curvature. In the embodiment illustrated in FIG. 1, nozzlesegment 451 is part of the first stage turbine nozzle adjacentcombustion chamber 390. In other embodiments, nozzle segment 451 islocated within a turbine nozzle 450 of another stage.

Pressure side wall 463 may span from leading edge 461 to trailing edge462 between upper endwall 453 and lower endwall 457. Pressure side wall463 may include a concave shape. Pressure side wall 463 may also includea pressure side surface 469, the outer surface of pressure side wall463, with a concave shape. Suction side wall 464 may also span fromleading edge 461 to trailing edge 462 between upper endwall 453 andlower endwall 457. Suction side wall 464 may include a convex shape.Leading edge 461, trailing edge 462, pressure side wall 463 and suctionside wall 464 may form a cooling cavity 485 (illustrated in FIGS. 3 and6) or cooling cavities there between. Upper endwall 453, lower endwall457, or both may include a hole, holes, or pathways for cooling air (notshown) to enter the cooling cavity 485.

Airfoil 460 may also include multiple groupings of film cooling holes orapertures. Each cooling hole or aperture may be a channel extendingthrough a wall of the airfoil, such as the pressure side wall 463. Inthe embodiment illustrated in FIG. 2, airfoil 460 includes showerheadcooling apertures 465, forward cooling apertures 466, aft coolingapertures 467, and intermediate cooling apertures 468. Showerheadcooling apertures 465 are located at leading edge 461 and may be groupedtogether along leading edge 461. Showerhead cooling apertures 465 may bearranged in columns. In the embodiment shown in FIG. 2, showerheadcooling apertures 465 are arranged in six columns, each column extendingin the radial direction between upper endwall 453 and lower endwall 457.In other embodiments, showerhead cooling apertures 465 may be arrangedin four to seven columns or may be arranged in other configurations. Theportions of pressure side wall 463 and suction side wall 464 adjacentleading edge 461 may include showerhead cooling apertures 465.

Forward cooling apertures 466 may be grouped together and located withinthe third of pressure side wall 463 that is adjacent leading edge 461.Forward cooling apertures 466 may be proximate showerhead coolingapertures 465. In embodiments, forward cooling apertures 466 are locatedfrom ⅛ to ¼ of the length of pressure side wall 463 from showerheadcooling apertures 465. In other embodiments, forward cooling apertures466 are located 1/6 of the length of pressure side wall 463 fromshowerhead cooling apertures 465. In yet other embodiments, forwardcooling apertures 466 are located at least ⅛ of the length of pressureside wall 463 from showerhead cooling apertures 465. Forward coolingapertures 466 may be grouped together between upper endwall 453 andlower endwall 457. In the embodiment illustrated in FIG. 2, forwardcooling apertures 466 are arranged in a single radial column and spacedapart radially at 3.5 pitch over diameter, the distance between thecenters of adjacent apertures over the diameter of the apertures. Inother embodiments, forward cooling apertures 466 are spaced apartradially from 3 to 4 pitch over diameter. Forward cooling apertures 466may overlap with an adjacent forward cooling aperture 466 rather thanalign in the radial direction along the surface of pressure side wall463.

Aft cooling apertures 467 may be grouped together and located within thethird of pressure side wall 463 that is adjacent to trailing edge 462.Aft cooling apertures 467 may be proximate trailing edge 462. Inembodiments, aft cooling apertures are located from ⅛ to ¼ of the lengthof pressure side wall 463 from trailing edge 462. In other embodiments,aft cooling apertures 467 are located 1/6 of the length of pressure sidewall 463 from trailing edge 462. In yet other embodiments, aft coolingapertures 467 are located at least ⅛ of the length of pressure side wall463 from trailing edge 462. Aft cooling apertures 467 may be arrangedradially between upper endwall 453 and lower endwall 457. In theembodiment illustrated in FIG. 2, aft cooling apertures 467 are arrangedin a single radial column and spaced apart radially at 3.5 pitch overdiameter. In other embodiments, aft cooling apertures 467 are spacedapart radially from 3 to 4 pitch over diameter. Aft cooling apertures467 may overlap with an adjacent aft cooling aperture 467 rather thanalign in the radial direction along the surface of pressure side wall463.

Intermediate cooling apertures 468 may be grouped together and locatedwithin the middle third of pressure side wall 463. Intermediate coolingapertures 468 may be between forward cooling apertures 466 and trailingedge 462. Intermediate cooling apertures 468 may also be between forwardcooling apertures 466 and aft cooling apertures 467. In someembodiments, intermediate cooling apertures 468 are located from ¼ to ⅜of the length of pressure side wall 463 from forward cooling apertures466 and 1/4 to 3/8 of the length of pressure side wall 463 from aftcooling apertures 467. In other embodiments, intermediate coolingapertures 468 are located 1/3 of the length of pressure side wall 463from forward cooling apertures 466 and 1/3 of the length of pressureside wall 463 from aft cooling apertures 467. In yet other embodiments,intermediate cooling apertures 468 are located at least ⅛ of the lengthof pressure side wall 463 from forward cooling apertures 466 and atleast ⅛ of the length of pressure side wall 463 from aft coolingapertures 467. Intermediate cooling apertures 468 may be arrangedradially between upper endwall 453 and lower endwall 457. In theembodiment illustrated in FIG. 2, intermediate cooling apertures 468 arearranged in a single radial column and spaced apart radially at 3.5pitch over diameter. In other embodiments, intermediate coolingapertures 468 are spaced apart radially from 3 to 4 pitch over diameter.Intermediate cooling apertures 468 may overlap with an adjacentintermediate cooling aperture 468 rather than being aligned in theradial direction along the surface of pressure side wall 463.

While the embodiment illustrated in FIG. 2 includes forward coolingapertures 466, aft cooling apertures 467, and intermediate coolingapertures 468, some embodiments do not include aft cooling apertures 467and other embodiments include second intermediate cooling apertureslocated between intermediate cooling apertures 468 and aft coolingapertures 467. The second intermediate cooling apertures may be arrangedsimilar to the arrangements of forward cooling apertures 466, aftcooling apertures 467, and intermediate cooling apertures 468. Othergroups or columns of cooling apertures may also be included. The spacingbetween groups or columns of cooling apertures may be dependent on thenumber of groups or columns of cooling apertures located along pressureside wall 463.

Airfoil 460 may further include slots 483. Slots 483 may be located onpressure side wall 463 and may be adjacent trailing edge 462. Slots 483may be rectangular and may be aligned in the radial direction betweenupper endwall 453 and lower endwall 457. Slots 483 may extend fromcooling cavity 485 to trailing edge 462.

In the embodiment illustrated in FIG. 2, nozzle segment 451 includessecond airfoil 470. Second airfoil 470 may include the same or similarfeatures as airfoil 460 including second leading edge 471, secondtrailing edge (not shown), second pressure side wall 473, and secondsuction side wall 474. Second airfoil 470 may further include secondshowerhead cooling apertures 475, second forward cooling apertures 476,second aft cooling apertures 477, second intermediate cooling apertures478, and second slots (not shown). The description of second leadingedge 471, the second trailing edge, second pressure side wall 473,second suction side wall 474, second showerhead cooling apertures 475,second forward cooling apertures 476, second aft cooling apertures 477,second intermediate cooling apertures 478, and the second slots may beoriented in the same or a similar manner as leading edge 461, trailingedge 462, pressure side wall 463, suction side wall 464, showerheadcooling apertures 465, forward cooling apertures 466, aft coolingapertures 467, intermediate cooling apertures 468, and slots 483respectively. In other embodiments, nozzle segment 451 only includesairfoil 460 and not second airfoil 470.

The various components of nozzle segment 451 including upper shroud 452,lower shroud 456, airfoil 460, and second airfoil 470 may be integrallycast or metalurgically bonded to form a unitary or one piece assemblythereof.

In accordance with embodiments of this invention, the spaced apartgroups of cooling apertures, showerhead cooling apertures 465, forwardcooling apertures 466, intermediate cooling apertures 468, and aftcooling apertures 467, alternate in directionality, being angled orpartially angled at lower endwall 457 or upper endwall 453. Thedirectionality or angle of the apertures directs cooling air in aselected direction. In the embodiment illustrated in FIGS. 2-6,showerhead cooling apertures 465 are angled toward lower endwall 457,forward cooling apertures 466, the next grouping of cooling holes alongpressure side wall 463, are angled toward upper endwall 453,intermediate cooling apertures 468, the following grouping of coolingholes along pressure side wall 463, are also angled at lower endwall457, and aft cooling apertures 467, the last grouping of cooling holes,are also angled at upper endwall 453. In other embodiments, showerheadcooling apertures 465 are angled toward upper endwall 453, forwardcooling apertures 466 are angled toward lower endwall 457, intermediatecooling apertures 468 are also angled at upper endwall 453, and aftcooling apertures 467 are also angled at lower endwall 457.

In embodiments that include the second intermediate cooling aperturesand showerhead cooling apertures 465 angled toward lower endwall 457,second intermediate cooling apertures would be the grouping afterintermediate cooling apertures 468 and would be angled towards upperendwall 453 and aft cooling apertures 467 would be the grouping afterthe second intermediate cooling apertures and would be angled towardlower endwall 457.

FIG. 3 is a cross-sectional view of a portion of the nozzle segment 451of FIG. 2 showing the showerhead cooling apertures 465. In theembodiment illustrated in FIG. 3, showerhead cooling apertures 465 mayextend through a wall 444 of airfoil 460 a cooling cavity 485 towardslower endwall 457. Wall 444 may be a part of leading edge 461, pressureside wall 463, or suction side wall 464. Showerhead cooling apertures465 may be angled relative to a reference plane 480. Reference plane 480may be defined as a plane perpendicular to a radial extending from thenozzle axis, the axis of upper shroud 452 and lower shroud 456 and alongleading edge 461. In one embodiment, showerhead angle 481, the angle ofshowerhead cooling apertures 465 relative to reference plane 480 is fromtwenty to forty-five degrees towards the lower endwall 457. In anotherembodiment, showerhead angle 481 is thirty degrees or within apredetermined tolerance of thirty degrees. The predetermined tolerancemay be the engineering tolerance or the manufacturing tolerance. Whileshowerhead angle 481 is directed toward lower endwall 457 in theembodiment illustrated, showerhead angle 481 may be directed towardupper endwall 453 in other embodiments.

Each showerhead cooling aperture 465 may also be angled towards thelower endwall 457 or the upper endwall 453 relative to the directionnormal to leading edge 461 at the location where the showerhead coolingaperture 465 is located.

As illustrated in FIG. 3, each showerhead cooling aperture 465 mayinclude showerhead inlet end 491 adjacent cooling cavity 485 andshowerhead outlet end 492 located at the surface of leading edge 461.Showerhead inlet end 491 may be radially closer to the upper endwall 453than showerhead outlet end 492 and showerhead outlet end 492 may beradially closer to the lower endwall 457 than showerhead inlet end 491.

FIG. 4 is a detailed view of the forward cooling apertures 466 of FIG.2. FIG. 5 is a detailed view of the intermediate cooling apertures 468of FIG. 2. FIG. 6 is a detailed view of the aft cooling apertures 467 ofFIG. 2. Referring to FIGS. 4, 5, and 6, forward cooling apertures 466,aft cooling apertures 467, and intermediate cooling apertures 468 may beangled relative to the flow direction of the air traveling throughturbine nozzle 450 along pressure side surface 469 during operation ofgas turbine engine 100. Reference line 482 illustrates the flowdirection. Reference line 482 may also be defined as the intersectionbetween pressure side surface 469 and a plane perpendicular to a radialextending from the turbine nozzle axis, the axis of upper shroud 452 andlower shroud 456, along the pressure side surface 469.

Referring to FIGS. 2 and 4, forward cooling apertures 466 may be angledat a forward compound angle 486. Forward compound angle 486 may be thecomponent of the angle of forward cooling apertures 466 in the plane ofpressure side surface 469. As illustrated, forward compound angle 486 isangled toward upper endwall 453 relative to the flow direction orreference line 482. In one embodiment, forward compound angle 486 isfrom fifteen to forty-five degrees. In another embodiment, forwardcompound angle 486 is thirty degrees or within a predetermined toleranceof thirty degrees. The predetermined tolerance may be the engineeringtolerance or the manufacturing tolerance. Zero degrees may be the flowdirection of the direction along reference line 482 traveling fromleading edge 461 to trailing edge 462. While forward compound angle 486is directed towards upper endwall 453 in the embodiment illustrated,forward compound angle 486 is directed towards lower endwall 457 inembodiments where showerhead angle 481 is directed towards upper endwall453.

Referring to FIGS. 2 and 5, intermediate cooling apertures 468 may beangled at an intermediate compound angle 488. Intermediate compoundangle 488 may be the component of the angle of intermediate coolingapertures 468 in the plane of pressure side surface 469. As illustrated,intermediate compound angle 488 is angled toward the lower endwall 457relative to the flow direction or reference line 482. In one embodiment,intermediate compound angle 488 is from fifteen to forty-five degrees.In another embodiment, intermediate compound angle 488 is thirty degreesor within a predetermined tolerance of thirty degrees. The predeterminedtolerance may be the engineering tolerance or the manufacturingtolerance. While intermediate compound angle 488 is directed towardslower endwall 457 in the embodiment illustrated, intermediate compoundangle 488 is directed towards upper endwall 453 in embodiments whereforward compound angle 486 is directed towards lower endwall 457.

Referring to FIGS. 2 and 6, aft cooling apertures 467 may be angled atan aft compound angle 487. Aft compound angle 487 may be the componentof the angle of aft cooling apertures 467 in the plane of pressure sidesurface 469. As illustrated, aft compound angle 487 is angled toward theupper endwall 453 relative to the flow direction or reference line 482and may be similar or equal to forward compound angle 486. In oneembodiment, aft compound angle 487 is from fifteen to forty-fivedegrees. In another embodiment, aft compound angle 487 is thirty degreesor within a predetermined tolerance of thirty degrees. The predeterminedtolerance may be the engineering tolerance or the manufacturingtolerance. While aft compound angle 487 is directed towards upperendwall 453 in the embodiment illustrated, aft compound angle 487 may bedirected towards lower endwall 457 in embodiments where intermediatecompound angle 488 is directed towards upper endwall 453.

In embodiments not including intermediate cooling apertures 468 orembodiments including second intermediate cooling apertures, aftcompound angle 487 may be angled toward lower endwall 457 relative tothe flow direction or reference line 482. In one of these embodiments,aft compound angle 487 is from fifteen to forty-five degrees. In anotherof these embodiments, aft compound angle 487 is approximately thirtydegrees.

FIG. 7 is a cross-section of the airfoil 460 of FIG. 2. Referring toFIG. 7, forward cooling apertures 466 may also include a forwardinjection angle 441. Forward injection angle 441 may be the component ofthe angle of forward cooling apertures 466 in the plane perpendicular topressure side surface 469. Forward injection angle 441 may be measuredrelative to a line extending toward trailing edge 462 and tangent topressure side surface 469 at the location of each forward coolingaperture 466. In one embodiment, forward injection angle 441 is fromfifteen to fifty degrees. In another embodiment, forward injection angle441 is approximately thirty degrees.

Aft cooling apertures 467 may also include an aft injection angle 442.Aft injection angle 442 may be the component of the angle of aft coolingapertures 467 in the plane perpendicular to pressure side surface 469.Aft injection angle 442 may be measured relative to a line extendingtoward trailing edge 462 and tangent to pressure side surface 469 at thelocation of each aft cooling aperture 467. In one embodiment, aftinjection angle 442 is from fifteen to fifty degrees. In anotherembodiment, aft injection angle 442 is approximately thirty degrees.

Intermediate cooling apertures 468 may also include an intermediateinjection angle 443. Intermediate injection angle 443 may be thecomponent of the angle of intermediate cooling apertures 468 in theplane perpendicular to pressure side surface 469. Intermediate injectionangle 443 may be measured relative to a line extending toward trailingedge 462 and tangent to pressure side surface 469 at the location ofeach intermediate cooling aperture 468. In one embodiment, intermediateinjection angle 443 is from fifteen to fifty degrees. In anotherembodiment, intermediate injection angle 443 is approximately thirtydegrees.

Cooling cavity 485 may be a single cavity or may be subdivided intomultiple cavities. In the embodiment illustrated in FIG. 7, coolingcavity 485 is subdivided into two cooling cavities.

Each forward cooling aperture 466 may include forward inlet end 493adjacent cooling cavity 485 and forward outlet end 494 adjacent or atpressure side surface 469. Each intermediate cooling aperture 468 mayinclude intermediate inlet end 497 adjacent cooling cavity 485 andintermediate outlet end 498 adjacent or at pressure side surface 469.Each aft cooling aperture 467 may include aft inlet end 495 adjacentcooling cavity 485 and aft outlet end 496 adjacent or at pressure sidesurface 469.

The compound angles may be determined by the positions of the inlet endsand the outlet ends of the apertures relative to lower endwall 457 andupper endwall 453, while the injection angle may be determined by thepositions of the inlet ends and the outlet ends relative to leading edge461 and trailing edge 462.

In the embodiment illustrated in FIG. 4, forward inlet end 493 isradially closer to lower endwall 457 and axially closer to leading edge461 than forward outlet end 494, and forward outlet end 494 is radiallycloser to upper endwall 453 and axially closer to trailing edge 462 thanforward inlet end 493. In other embodiments, forward inlet end 493 isradially closer to upper endwall 453 and axially closer to leading edge461 than forward outlet end 494, and forward outlet end 494 is radiallycloser to lower endwall 457 and axially closer to trailing edge 462 thanforward inlet end 493.

In the embodiment illustrated in FIG. 5, intermediate inlet end 497 isradially closer to upper endwall 453 and axially closer to leading edge461 than intermediate outlet end 498, and intermediate outlet end 498 isradially closer to lower endwall 457 and axially closer to trailing edge462 than intermediate inlet end 497. In other embodiments, intermediateinlet end 497 is radially closer to lower endwall 457 and axially closerto leading edge 461 than intermediate outlet end 498, and intermediateoutlet end 498 is radially closer to upper endwall 453 and axiallycloser to trailing edge 462 than intermediate inlet end 497.

In the embodiment illustrated in FIG. 6, aft inlet end 495 is radiallycloser to lower endwall 457 and axially closer to leading edge 461 thanaft outlet end 496, and aft outlet end 496 is radially closer to upperendwall 453 and axially closer to trailing edge 462 than aft inlet end495. In other embodiments, aft inlet end 495 is radially closer to upperendwall 453 and axially closer to leading edge 461 than aft outlet end496, and aft outlet end 496 is radially closer to lower endwall 457 andaxially closer to trailing edge 462 than aft inlet end 495.

One or more of the above components (or their subcomponents) may be madefrom stainless steel and/or durable, high temperature materials known as“superalloys”. A superalloy, or high-performance alloy, is an alloy thatexhibits excellent mechanical strength and creep resistance at hightemperatures, good surface stability, and corrosion and oxidationresistance. Superalloys may include materials such as HASTELLOY, alloyx, INCONEL, WASPALOY, RENE alloys, HAYNES alloys, alloy 188, alloy 230,INCOLOY, MP98T, TMS alloys, and CMSX single crystal alloys.

INDUSTRIAL APPLICABILITY

Gas turbine engines may be suited for any number of industrialapplications such as various aspects of the oil and gas industry(including transmission, gathering, storage, withdrawal, and lifting ofoil and natural gas), the power generation industry, cogeneration,aerospace, and other transportation industries.

Referring to FIG. 1, a gas (typically air 10) enters the inlet 110 as a“working fluid”, and is compressed by the compressor 200. In thecompressor 200, the working fluid is compressed in an annular flow path115 by the series of compressor disk assemblies 220. In particular, theair 10 is compressed in numbered “stages”, the stages being associatedwith each compressor disk assembly 220. For example, “4th stage air” maybe associated with the 4th compressor disk assembly 220 in thedownstream or “aft” direction, going from the inlet 110 towards theexhaust 500). Likewise, each turbine disk assembly 420 may be associatedwith a numbered stage.

Once compressed air 10 leaves the compressor 200, it enters thecombustor 300, where it is diffused and fuel is added. Air 10 and fuelare injected into the combustion chamber 390 via fuel injector 310 andcombusted. Energy is extracted from the combustion reaction via theturbine 400 by each stage of the series of turbine disk assemblies 420.Exhaust gas 90 may then be diffused in exhaust diffuser 520, collectedand redirected. Exhaust gas 90 exits the system via an exhaust collector550 and may be further processed (e.g., to reduce harmful emissions,and/or to recover heat from the exhaust gas 90).

Operating efficiency of a gas turbine engine generally increases with ahigher combustion temperature. Thus, there is a trend in gas turbineengines to increase the temperatures. Gas reaching forward stages of aturbine from a combustion chamber 390 may be 1000 degrees Fahrenheit ormore. To operate at such high temperatures a portion of the compressedair from the compressor 200, cooling air, may be diverted throughinternal passages or chambers to cool various components of a turbineincluding turbine nozzle segments such as nozzle segment 451. However,the use of cooling air may reduce the operating efficiency of the gasturbine engine.

Alternating the direction of groupings of cooling apertures such asshowerhead cooling apertures 465, forward cooling apertures 466,intermediate cooling apertures 468, and aft cooling apertures 467, todirect cooling air towards upper endwall 453 of upper shroud 452 andlower endwall 457 of lower shroud 456 may reduce the temperatures ofupper endwall 453 and lower endwall 457, which may improve the operatinglife of nozzle segment 451.

The first order cooling or initial use of the cooling air exitingshowerhead cooling apertures 465, forward cooling apertures 466,intermediate cooling apertures 468, and aft cooling apertures 467 may beto film cool pressure side wall 463. The second order cooling or seconduse of the cooling air may be to reduce the temperatures of upperendwall 453 and lower endwall 457.

The cooling air may be directed through turbine housing 430, turbinediaphragm 440, or both and into cooling cavity 485. The cooling air maythen be directed through the cooling apertures including showerheadcooling apertures 465, forward cooling apertures 466, intermediatecooling apertures 468, and aft cooling apertures 467. The cooling airmay also be used for cooling airfoil 460 internally prior to passingthrough the cooling apertures. The multiple uses of the cooling air thatmay include the first order film cooling, the second order endwallcooling, and the internal cooling may reduce the amount of air needed toeffectively cool nozzle segment 451. Reducing the amount of air neededto cool nozzle segment 451 may improve or increase the efficiency of gasturbine engine 100.

The cooling apertures of second airfoil 470 may be used in the same or asimilar manner as the cooling apertures of airfoil 460 resulting in afurther reduction of the temperatures of upper endwall 453 and lowerendwall 457, as well as the reduction in the amount of cooling airneeded to effectively cool each nozzle segment 451.

The preceding detailed description is merely exemplary in nature and isnot intended to limit the invention or the application and uses of theinvention. The described embodiments are not limited to use inconjunction with a particular type of gas turbine engine. Hence,although the present disclosure, for convenience of explanation, depictsand describes a particular nozzle segment, it will be appreciated thatthe nozzle segment in accordance with this disclosure can be implementedin various other configurations, can be used with various other types ofgas turbine engines, and can be used in other types of machines.Furthermore, there is no intention to be bound by any theory presentedin the preceding background or detailed description. It is alsounderstood that the illustrations may include exaggerated dimensions tobetter illustrate the referenced items shown, and are not considerlimiting unless expressly stated as such.

What is claimed is:
 1. A nozzle segment for a nozzle ring of a gasturbine engine, the nozzle segment comprising: a first endwall; a secondendwall; and an airfoil extending between the first endwall and thesecond endwall, the airfoil including a leading edge extending radiallyfrom the first endwall to the second endwall, a trailing edge extendingradially from the first endwall to the second endwall axially distal tothe leading edge, a pressure side wall extending from the leading edgeto the trailing edge, a suction side wall extending from the leadingedge to the trailing edge, a plurality of showerhead cooling aperturesspanning along the leading edge, a plurality of forward coolingapertures grouped together in the pressure side wall proximate theplurality of showerhead cooling apertures, and a plurality ofintermediate cooling apertures grouped together in the pressure sidewall between the trailing edge and the plurality of forward coolingapertures, the plurality of showerhead cooling apertures, the pluralityof forward cooling apertures, and the plurality of intermediate coolingapertures alternating in directionality such that the plurality ofshowerhead cooling apertures are angled toward the first endwall, theplurality of forward cooling apertures are angled toward the secondendwall, and the plurality of intermediate cooling apertures are angledtoward the first endwall.
 2. The nozzle segment of claim 1, wherein eachshowerhead cooling aperture is angled toward the first endwall as eachshowerhead cooling aperture extends through a wall of the airfoil, eachforward cooling aperture includes a forward compound angle from fifteento forty-five degrees towards the second endwall relative to a flowdirection of air through the nozzle segment during operation, and eachintermediate cooling aperture including an intermediate compound anglefrom fifteen to forty-five degrees towards the first endwall relative toa flow direction of air through the nozzle segment during operation. 3.The nozzle segment of claim 2, wherein the airfoil further includes aplurality of aft cooling apertures grouped together in the pressure sidewall proximate the trailing edge, each aft cooling aperture including anaft compound angle from fifteen to forty-five degrees towards the secondendwall relative to a flow direction of air through the nozzle segmentduring operation.
 4. The nozzle segment of claim 1, wherein theplurality of forward cooling apertures are arranged in a single columnand the plurality of intermediate cooling apertures are arranged in asingle column.
 5. The nozzle segment of claim 1, wherein the firstendwall is a lower endwall and the second endwall is an upper endwalllocated radially outward from the lower endwall.
 6. The nozzle segmentof claim 1, wherein each forward cooling aperture is spaced apart froman adjacent forward cooling aperture from 3 to 4 pitch over diameter andeach intermediate cooling aperture is spaced apart from an adjacentintermediate cooling aperture from 3 to 4 pitch over diameter.
 7. Thenozzle segment of claim 1, wherein the plurality of showerhead coolingapertures is configured to direct air to film cool the leading edge andcool the first endwall, the plurality of forward cooling apertures isconfigured to direct air to film cool a pressure side surface of thepressure side wall and cool the second endwall, and the plurality ofintermediate cooling apertures is configured to direct air to film coolthe pressure side surface and cool the first endwall.
 8. A gas turbineengine including the nozzle segment of claim 1, wherein the nozzlesegment is located in a first stage turbine nozzle of the gas turbineengine.
 9. A nozzle segment for a nozzle ring of a gas turbine engine,the nozzle segment comprising: a first endwall; a second endwall; anairfoil extending between the first endwall and the second endwall, theairfoil including a leading edge extending radially from the firstendwall to the second endwall, a trailing edge extending radially fromthe first endwall to the second endwall, a pressure side wall extendingfrom the leading edge to the trailing edge, the pressure side wallincluding a pressure side surface with a concave shape, the pressureside surface being the outer surface of the pressure side wall, asuction side wall extending from the leading edge to the trailing edge,a cooling cavity located between the leading edge, the trailing edge,the pressure side wall, and the suction side wall, a plurality ofshowerhead cooling apertures spanning along the leading edge, eachshowerhead cooling aperture including a showerhead inlet end adjacentthe cooling cavity and a showerhead outlet end at the outer surface ofthe leading edge, the showerhead inlet end being radially closer to thesecond endwall than the showerhead outlet end and the showerhead outletend being radially closer to the first endwall than the showerhead inletend, a plurality of forward cooling apertures in the pressure side wallgrouped together and spaced apart from the plurality of showerheadcooling apertures at least ⅛ the length of the pressure side wall, eachforward cooling aperture including a forward inlet end adjacent thecooling cavity and a forward outlet end adjacent the pressure sidesurface, the forward inlet end being radially closer to the firstendwall and axially closer to the leading edge than the forward outletend, and the forward outlet end being radially closer to the secondendwall and axially closer to the trailing edge than the forward inletend, and a plurality of intermediate cooling apertures in the pressureside wall grouped together and spaced apart from the plurality offorward cooling apertures at least ⅛ the length of the pressure sidewall, each intermediate cooling aperture including an intermediate inletend adjacent the cooling cavity and an intermediate outlet end adjacentthe pressure side surface, the intermediate inlet end being radiallycloser to the second endwall and axially closer to the leading edge thanthe intermediate outlet end, and the intermediate outlet end beingradially closer to the first endwall and axially closer to the trailingedge than the intermediate inlet end.
 10. The nozzle segment of claim 9,wherein the airfoil further includes a plurality of aft coolingapertures in the pressure side wall grouped together and spaced apartfrom the plurality of intermediate cooling apertures at least ⅛ thelength of the pressure side wall, each aft cooling aperture including anaft inlet end adjacent the cooling cavity and an aft outlet end adjacentthe pressure side surface, the aft inlet end being radially closer tothe first endwall and axially closer to the leading edge than the aftoutlet end, and the aft outlet end being radially closer to the secondendwall and axially closer to the trailing edge than the aft inlet end.11. The nozzle segment of claim 9, wherein the first endwall is a lowerendwall of a lower shroud and the second endwall is an upper endwall ofan upper shroud, the lower endwall being located radially inward fromthe upper endwall.
 12. The nozzle segment of claim 9, wherein eachforward cooling aperture is spaced apart from an adjacent forwardcooling aperture from 3 to 4 pitch over diameter and each intermediatecooling aperture is spaced apart from an adjacent intermediate coolingaperture from 3 to 4 pitch over diameter.
 13. The nozzle segment ofclaim 10, wherein each forward cooling aperture is spaced apart from anadjacent forward cooling aperture from 3 to 4 pitch over diameter, eachintermediate cooling aperture is spaced apart from an adjacentintermediate cooling aperture from 3 to 4 pitch over diameter, and eachaft cooling aperture is spaced apart from an adjacent aft coolingaperture from 3 to 4 pitch over diameter.
 14. A nozzle segment for anozzle ring of a gas turbine engine, the nozzle segment comprising: anupper shroud including an upper endwall, the upper endwall being theshape of a sector of a toroid; a lower shroud including a lower endwalllocated radially inward from the upper endwall, the lower endwall beingthe shape of a sector of a toroid; an airfoil extending between theupper endwall and the lower endwall, the airfoil including a leadingedge extending radially from the upper endwall to the lower endwall, atrailing edge extending radially from the upper endwall to the lowerendwall axially distal to the leading edge, a pressure side wallextending from the leading edge to the trailing edge, the pressure sidewall including a concave shape, a suction side wall extending from theleading edge to the trailing edge, the suction side wall including aconvex shape, a cooling cavity located between the leading edge, thetrailing edge, the pressure side wall, and the suction side wall, aplurality of showerhead cooling apertures arranged in four to sevencolumns spanning along the leading edge, each showerhead coolingaperture being angled toward the lower endwall as each showerheadcooling aperture extends through a wall of the airfoil from the coolingcavity, a plurality of forward cooling apertures arranged in a columnextending radially between the upper endwall and the lower endwall andlocated in the third of the pressure side wall adjacent the leadingedge, each forward cooling aperture extending through the pressure sidewall from the cooling cavity and including a forward compound angle fromfifteen to forty-five degrees towards the upper endwall and the trailingedge relative to a reference line in the plane of a pressure sidesurface of the pressure side wall, the reference line being defined asan intersection between the pressure side surface and a planeperpendicular to a radial extending from an axis of the upper shroudalong the pressure side surface, a plurality of intermediate coolingapertures arranged in a column extending radially between the upperendwall and the lower endwall and located in the middle third of thepressure side wall between the leading edge and the trailing edge, eachintermediate cooling aperture extending through the pressure side wallfrom the cooling cavity and including an intermediate compound anglefrom fifteen to forty-five degrees towards the lower endwall and thetrailing edge relative to the reference line, and a plurality of aftcooling apertures arranged in a column extending radially between theupper endwall and the lower endwall and located in the third of thepressure side wall adjacent the trailing edge, each aft cooling apertureextending through the pressure side wall from the cooling cavity andincluding an aft compound angle from fifteen to forty-five degreestowards the upper endwall and the trailing edge relative to thereference line.
 15. The nozzle segment of claim 14, further comprising:a second airfoil extending between the upper endwall and the lowerendwall, the second airfoil including a second leading edge extendingradially from the upper endwall to the lower endwall, a second trailingedge extending radially from the upper endwall to the lower endwalldistal to the second leading edge, a second pressure side wall extendingfrom the second leading edge to the second trailing edge, the secondpressure side wall including a second concave shape, a second suctionside wall extending from the second leading edge to the second trailingedge, the second suction side wall including a second convex shape, asecond cooling cavity located between the second leading edge, thesecond trailing edge, the second pressure side wall, and the secondsuction side wall, a plurality of second showerhead cooling aperturesarranged in four to seven columns spanning along the second leadingedge, each second showerhead cooling aperture being angled toward thelower endwall as each second showerhead cooling aperture extends througha second wall of the second airfoil from one of the second coolingcavity, a plurality of second forward cooling apertures arranged in acolumn extending radially between the upper endwall and the lowerendwall and located in the third of the second pressure side walladjacent the second leading edge, each second forward cooling apertureextending through the second pressure side wall from the second coolingcavity and including a second forward compound angle from fifteen toforty-five degrees towards the upper endwall and the second trailingedge relative to a second reference line in the plane of a secondpressure side surface of the second pressure side wall, the secondreference line being defined as an intersection between the secondpressure side surface and a second plane perpendicular to a radialextending from the axis of the upper shroud along the second pressureside surface, a plurality of second intermediate cooling aperturesarranged in a column extending radially between the upper endwall andthe lower endwall and located in the middle third of the second pressureside wall between the second leading edge and the second trailing edge,each second intermediate cooling aperture extending through the secondpressure side wall from the second cooling cavity and including a secondintermediate compound angle from fifteen to forty-five degrees towardsthe lower endwall and the second trailing edge relative to the secondreference line, and a plurality of second aft cooling apertures arrangedin a column extending radially between the upper endwall and the lowerendwall and located in the third of the second pressure side walladjacent the second trailing edge, each second aft cooling apertureextending through the second pressure side wall from the second coolingcavity and including a second aft compound angle from fifteen toforty-five degrees towards the upper endwall and the second trailingedge relative to the second reference line.
 16. The nozzle segment ofclaim 14, wherein each forward cooling aperture is spaced apart from anadjacent forward cooling aperture from 3 to 4 pitch over diameter, eachintermediate cooling aperture is spaced apart from an adjacentintermediate cooling aperture from 3 to 4 pitch over diameter, and eachaft cooling aperture is spaced apart from an adjacent aft coolingaperture from 3 to 4 pitch over diameter.
 17. The nozzle segment ofclaim 14, wherein each forward cooling aperture is spaced apart from anadjacent forward cooling aperture at 3.5 pitch over diameter, eachintermediate cooling aperture is spaced apart from an adjacentintermediate cooling aperture at 3.5 pitch over diameter, and each aftcooling aperture is spaced apart from an adjacent aft cooling apertureat 3.5 pitch over diameter.
 18. The nozzle segment of claim 14, whereinthe plurality of forward cooling apertures are spaced apart from theplurality of showerhead cooling apertures from ⅛ to ¼ the length of thepressure side wall, the plurality of intermediate cooling apertures arespaced apart from the plurality of forward cooling apertures from ¼ to ⅜the length of the pressure side wall, and the plurality of aft coolingapertures are spaced apart from the plurality of intermediate coolingapertures from ¼ to ⅜ of the length of the pressure side wall.
 19. Thenozzle segment of claim 14, wherein the forward compound angle is withina predetermined tolerance of thirty degrees, the intermediate compoundangle is within a predetermined tolerance of thirty degrees, and the aftcompound angle is within a predetermined tolerance of thirty degrees.20. A gas turbine engine including the nozzle segment of claim 14,wherein the nozzle segment is located in a first stage turbine nozzle ofthe gas turbine engine.